DE2830502C3 - Missile control device - Google Patents

Description

2. Control device according to claim 1, characterized in that the logic circuit contains a multiplier circuit (22) with four inputs (20,28,30,32), the first input (20) of which the output signal of the threshold switch (18) and the second input ( 32) a signal proportional to the time derivative (Ri) of the distance signal is fed from the proximity sensor (24), and the third and fourth inputs (28, 30) of which are connected to the output of a reversing circuit (26) which is acted upon by the distance signal (Ri) , and that the output signal (άι · κ) of the multiplying circuit (22) at the input (14) of the steering controller (12) the curve speed signal (ύή is switched off.

3. Control device according to claim 1, characterized in that the logic circuit reveals a distance threshold value switch (56) to which the proximity sensor (54) is acted upon by the distance signal (Ri) and when a distance value (Rn ) a switching signal (Y) ran and that a switch (40) is controlled by the switching signal (Y) , which switches off the line-of-sight rotational speed signal (ar) from the steering controller (48) and switches on the signal of the viewing angle threshold value switch (42) instead, whereby the signal of the squint angle | -Schwe | lwertschaIters corresponds to the double positive or negative limit value of the line of sight rotational speed (arc) , which results in the maximum final rudder torque.

The invention relates to a control device for missiles according to the preamble of the patent

Ii saying 1.

In a known control device (DE-PS 14 06 578) the homing head contains a gyro rotor, which carries an optical system designed as a Cassegrain system, the optical axis of which revolves with the axis of the gyro rotor coincides in the image plane of the optical system is also with the Rotary rotor rotating interrupter disc provided on one half z. B. alternately translucent and opaque sectors or a chess

2> board pattern and on the other half a system of having concentric, alternating translucent and opaque rings die through the Radiation passing through the glass and an infrared-permeable filter falls on you

jo infrared-sensitive receiver, which when detecting a target off the optical axis of the system emits a modulated AC signal. The phase of the alternating current signal obtained after demodulation depends on the direction of the target deviation

j-, on the optical axis of the system, and the The amplitude of the AC signal depends on the size of the target deposit. The gyro rotor is diagonal magnetized and surrounded by a ring coil. This is amplified accordingly on this ring coil

• κι alternating current signal switched, whereby periodically with the rotational frequency moments are exerted on the radially magnetized gyro rotor and this with the The axis of the optical system precessed towards the target. As a result, the gyro rotor with its orbital

r. axis and the axis of the optical system are constantly tracking the target.

The signals required for this in the ring coil are proportional to the gyroscopic laws Angular velocity of the line of sight in space, the

M, inertial line of sight velocity. Since a collision course is steered when the missile and target move in a straight line, if the line of sight from the missile to the target remains fixed in space, the signals given to the ring spool are as

γ, line of sight change signal c simultaneously given to the steering controller of the missile, so that the course change speed of the missile is proportional to the inertial curve speed of rotation.
Such missiles are used, for example, as aerial

wi used air missiles, taking the target from the hot Aircraft jet engine gases.

It is also known to measure the deflection of the axis of the Krciselrolors relative to the missile

h "i the longitudinal axis to be scanned and the" squint angle "closed determine. The so obtained Schiclwinkel signal is, however, in the known .Steuervorrichtung as sole steering command switched to the steering controller

tet, so that the squint angle is regulated to zero and the longitudinal axis of the missile aligned with the target will. In this way, however, the missile is guided to the target in a less favorable manner.

It is still one by DE-OS 19 51 518 Control device for missiles with target seeker known, in which a sensor for the "Squint angle" is provided between the missile longitudinal axis and the optical axis of the homing head as well a threshold switch via which an output signal dependent on the squint angle is sent to the steering controller is switchable.

In this known arrangement is to be prevented that with strong control commands one large transverse acceleration of the missile and a correspondingly large angle of attack require the Do not move the seeker head to the stop and you can lose the target as a result. For this reason then if the control command exceeds a predetermined threshold value and so does the squint angle is large, an intervention in the control is made in the form that a "and rrtit" dependent on the squint angle this rising signal is switched against the line of sight change signal at the input of the steering controller will. This does not mean that the lateral acceleration required for purely proportional navigation is commanded, but it is ensured that the seeker head is at a sufficient distance from the stop in his Work area remains and continues to focus on the goal

The threshold value switch thus speaks to a threshold value of the rate of change of the line of sight certain control signal and switches a function of the squint angle as a counter command on the input of the steering controller.

The hit storage of air-to-air missiles whose homing heads with infrared-sensitive receivers, InSb detectors, in particular, can be severely impaired in the vicinity of the target. It has showed that this is due to the presence of so-called "hot spots". Are “hot spots” Nodal areas in the exhaust jet of jet engines of modern combat aircraft, which are particularly important Emit radiation in the sensitivity range of InSb detectors. In afterburner operation, the Radiation intensity in the "Hol Spot *" can assume such high values that the hot spots of the target seeker head perceived and pursued as a phantom target. Since such hot spots are up to 15 m away from the Tail of the aircraft can occur, this means even with otherwise error-free steering of the missile a large hit list from the actual target.

But even if there is no disruption of target tracking by hot spots, /. IJ. for HF target seekers, it is generally desirable to direct the missile to a point that is opposite the engines somewhat in the direction of the aircraft longitudinal axis forward, e.g. B. towards the cockpit. offset is where a hit is usually more effective.

The invention is based on the object in a control device of the type defined at the outset Missile on one point to steer eggs opposite the point tracked by the homing head is offset forward in the direction of the longitudinal axis of the aircraft.

According to the invention, this object is achieved by the features listed in the characterizing part of claim 1 Measures resolved.

In the idealized I all linearly more uniform

ΙΊ

Movements of the target and missile, the missile is controlled in such a way that the line of sight to the target remains fixed in space. The missile longitudinal axis will adjust itself at a greater angle to the line of sight, the greater the angle between the line of sight and the aircraft speed vector. If the speed vector of the target aircraft runs in the direction of the line of sight, the missile must fly in the direction of the line of sight and aligns itself with its longitudinal axis on this and thus on the axis of the seeker head. The "squint angle" is then zero. However, if the aircraft moves at an angle transverse to the line of sight, the missile longitudinal axis must adjust itself at such an angle to the line of sight that it and thus also the direction of movement of the missile is directed to a point in the direction of movement of the aircraft in front of this point. A lead signal is only generated when the angle between the line of sight and the airspeed vector, and thus the squint angle, exceeds a certain predetermined amount. That is, the sign of the lead signal is determined by the sign of the squint angle.

The lead signal is also dependent on the distance between the missile and the target aircraft. In large distance of the missile from the target aircraft, no or practically no lead signal is required because Seen from a great distance the Winkelgeschwindig speed of the line of sight vector to the target aircraft and z. B. that of the line of sight vector to a "hot spot" practically coincide. A lead signal is effective at small distances from the target.

DE-AS 19 12 704 describes a method for self-steering one according to the radar target search method military missile heading for its target when it is used against sea targets known at which the steering elements of the missile can be adjusted in azimuth and elevation by radar error signals. Around not only to hit the parts of the sea target object protruding from the water and detected by the radar, it is provided there that an additional control voltage is generated from the radar error signals in the elevation plane is obtained, which influences the flight direction of the missile in the elevation plane in such a way that the missile hits its target object in the elevation plane at a predetermined distance below the point of impact that would have been taken without switching on the additional control voltage. The purpose of this measure is to build a ship not in the area of the deck superstructures but in the area of the waterline or below. To the The subject of DEAS 19 12 704 is based on a similar problem as the present application. It However, it is a different application and also υ; «-indere solvents.

In particular, in the arrangement according to the invention, a Vorh.i'ft in the direction of the - variable - Aircraft speed vector specified and not as in the known arrangement only one Distraction in the Klevationscbeno downwards.

Further refinements of the invention are the subject matter the utility claimC;

The invention is illustrated below in two exemplary embodiments explained in more detail with reference to the drawings.

Figure I illustrates the flight path geometry;

F i g. Fig. 2 is a block diagram of an embodiment of a control device according to the invention;

F i g. 3 shows the result of a simulation of FIG

Trajectories of a target aircraft and one with a missile controlled according to the invention;

Fig. Figure 4 is a block diagram of one opposite to that of Figure 4 Implementation form of FIG. 2 simplified control device.

Let us assume that the aircraft chasing / ■ / fin a distance Λ behind its tricbwcrkiiuslaB in the extension of the flight / .ciiglängsachsc a "hot spot" // 7SA *, and that the seeker head of the missile I'K stalls Aircraft engine followed this I lot spot. The missile IK would then be controlled according to the Ciesct / proportional navigation so that a change in the line of sight winkeis ii / '/ ii this I lot spot //7'.SViIH space is counteracted and the missile takes an undesirable collision course with the I lot Spot HTSP tracked. From the trajectory geometry according to I ι g. 1 results in the following relationship / wischen eiern Sichiiinienwinkci »r to the decoy target Hol Spot //7'.SViIiKl the line of sight angle α ι to the true target aircraft IV. And angular velocities between the corresponding lines of sight:

Fs is

sinl «,. .>,)

whereby

; ■ / the angle between the aircraft speed vector Vi and the reference direction W /: 7-'untl

Ri is the distance between the missile IK and the aircraft //.

Differentiation according to the line results in:

■:,

R ₁ cosl ·;,.

So a pre-echo signal must be switched on the steering controller to the o /.- Stciiersignal supplied by the seeker head, in order not to regulate after or but ii / and to steer the missile stall on the hot spot HTS1 ' on the aircraft.

Equation (2) can be simplified as follows:

1. The size

COS (rTp - .T,>

can be set approximately equal to one. At a greater distance from target isl n _r substantially or- equal to the angle differ to any significant extent only when such approach to the target, that then a correction of the steering command by a cosine function of the angle difference anyway no appreciable effect on the flight path of the missile FK shark.

2. In conventional air-to-air missiles, on the one hand, due to the high gain in the steering loop, (»/ ■ is small. The direction of flight of the target also changes relatively slowly, so that the term όι · -γι is relatively small. R /, ie the approach speed of the missile to the aircraft is relatively high, so that RiIRi becomes large compared to (> / ■ -} '; and the second term in brackets in equation (2) always outweighs the first term if the angular difference (i / · -) ¹ / does not get too small
The angle (> / ■ -).. '/ Between the line of sight to decoy //7".SVund the llugzcuggeschwindigkcit hünfi as already explained, with the angle of deviation A between Fliigkörpcrachsc and Zielsuchkopfiichsc together Fs, the sine function can of or-yi approximately · by a step function of the squint angle λ which assumes the value zero for small squint angles and above a critical angle λ λ ,, a fixed positive value K and within a negative ν irenzwinkeis -λ ,, a corresponding ncgaiivcn value - K.

This simplifies equation (2) / ii

■ ;, .f _r „ _rk for; ■ /.,; ■■ / - 'i _r for / </.,.

K ' sign (/. (
"ι"

The schiclwinkel λ remains during the whole If flight is small, it is essentially about a direct follow-up shot, so that the haul spots do not interfere with the steering of the missile. Ils in this case there is no correction signal »/ ·». aufgeschallel. A correction signal is activated for larger squint angles that occur when the aircraft moves across the line of sight.

The approach velocity Ri and the distance Ri of the missile to the target are required as essential parameters in equation (4). Both values are measured with the help of a proximity sensor. The distance A between Hol Spot HTSP and Fhigzcugheck is assumed to be constant and, together with the appropriately selected variable K, defines the gain of the correction.

(! ine circuit arrangement through which a correction of the control signal according to equations (3) and (4) is shown in FIG. 2 shown as a block diagram.

10 denotes the seeker head, which can be designed, for example, in the manner of DF-PS 14 06 578 and supplies a signal proportional to or , the angular velocity of the Sichllinienvcklors / u the target pursued by the seeker JO, which in the present case is the Hot Spot HTSP isl. This signal ä _r acts on the steering controller 12, the affected via an actuator 14, the F'lugkörperbcwcgung in a conventional manner. The missile movement and path kinematics are shown in FIG. 2 symbolized by the box 16. It results in an influence on Ru Ri and o / ·. This is the steering loop of a missile in proportional navigation.

The seeker head 10 also has a tap for the squint angle z. on. The signal supplied by this tap acts on a threshold switch iS. The Schweüwenschaiier ίο supplies the output signal zero for -A ₍ , <z. < + Ä, " For squint angle λ> λ ₍ . The threshold switch 18 ran a signal + AK.

for squint angle λ < -λα the output signal is equal to -A ■ K.

The output signal of the threshold switch 18 is at an input 20 of a multiplier circuit 22. A proximity sensor 24 supplies signals in accordance with the distance Rt between the missile and the target aircraft and the change in distance Rt- The distance signal Rt lifcg *. at a reciprocal circuit 26. This is practically a quotient generator, at whose counter input there is a constant signal "one" and whose denominator input receives the signal Ri . The output signal I // r of the circuit 26 is applied to two inputs 28, 30 of the multiplier circuit 22 at the same time. At a further input 32 of the Mulliplier circuit 22 is the signal Ar from the proximity sensor 24. The multiplier circuit 22 generates a signal proportional to the product of all signals present at the four inputs 20, 28, 30 and 32, and this signal is used as a lead signal on the basis of the steering controller ¹ - \ i the line of sight rotational speed signal r>/> aiifge · switched.

If the squint angle between -. / ,, and A + ₁ is located, the signal at the input 20 is zero, and accordingly is also the Ausgangssigmil of the multiplier 22 is zero. The missile FK is steered in the usual way according to the law of proportional navigation towards the hot spot. but which, viewed from the missile, is essentially in alignment with the actual target, so that the missile ultimately also hits the target. If the squint angle λ is greater than + Kc, or less than - '/.,. so the multiplier circuit delivers a lead signal άρκ according to equation (4). which is applied to the signal or in accordance with equation (3) at 34.

F i g. 3 shows the result of a simulation of the correction method described. It is about with a target approach of a realistic air-to-air missile. to a non-maneuvering target with the following initial conditions:

(- 0 I ι, = 300 m see launch speed of the missile *.
V ₁ - 3tK) m see speed of the

Target.

, τ ρ, = 60 angle of attack.

R _rn = 800 m firing distance.

After a flight time of 3.26 seconds, the in F i g. 3 set trajectory geometry shown. Up to this point in time, the missile was flying according to purely proportional navigation and reduced its distance to the target to 130 m. This flight condition is the starting point for the target approach with and without correction. The dashed line represents the flight path of the missile without correction. The cross connection between the missile and target path are the lines of sight between the missile seeker head and the hot spot at time intervals of 0.05 seconds. After 3.6 seconds, the missile flies past the HTSP hot spot with a deposit of approx. 0.2 m. If one assumes that the permissible hit deposition is 10 m, which is shown by the dashed circle, then the missile just no longer fulfills the hit condition.

With the correction described, the missile flies along the extended flight path. The missile crosses the trajectory 10 m in front of the hot spot and thus hits the target exactly. The hit condition at this moment is shown by the solid circle.

The controller can be further simplified when it is ensured that the approach speed Rr does not fall below a certain minimum value Rrmm, by

^AK R ² '< rc.

(5)

is defined.

In this inequality, <ij> ₍ . Means the limit

r> the line of sight turning speed. of the maximum Rudder torque of the missile. The steering controller receives one of the line of sight turning speed ο / · proportional signal and causes one of the Change of the line of sight counteracting · »RuüeiVieii-

. "i moment, which in turn is proportional to the speed of rotation of the line of sight. The design-related maximum rudder adjustment torque corresponds to a certain limit value on; the speed of rotation of the line of sight. Rh; is a limit value for the distance from the target, the

. ·. is determined by the fact that the design-related maximum possible transverse acceleration of the missile from this distance is a predetermined one optimal lead of the missile can be achieved.

When the given in equation (5) is met

ν condition, the missile can first be steered in the usual way according to the law of proportional navigation until the distance Rn from the target is reached. At this distance, the steering loop is broken and the steering controller is activated

i. Steering command n _n , activated, which effects the maximum lateral acceleration of the missile. This steering command η pe, then, because of the appropriate choice of the distance Rn; the desired reserve against the lip information received last

ι "tion specific course.

However, as with the Sipuervorrichtune from F i B. 2. only if the squint angle has given limit values ± /.,. exceeds. The signals supplied to the steering controller are

i "then

<T ₇ - .T _p for

and R _T <R _TC , or R _T > R _{T (} ,

π j = n _rG sign! /.) for >. >>. _o and R ₁ <R ,,,

A control device operating according to this principle is shown in FIG. 4 schematically as a block diagram shown

In Fig. 4, a seeker head is designated by 36, which supplies a signal öp proportional to the speed of rotation of the line of sight opinion and a signal proportional to the squint angle λ. The first signal is applied to a fixed contact 38 of a changeover switch 40 (which is shown here as a mechanical changeover switch but can be formed by an electronic component). The squint angle signal is applied to a threshold value switch 42 which, at a squint angle

- rc, <'■ < +> -G
an output signal zero at a squint angle λ> λα

an output signal + 2On; and at a squint angle λ <-Ar; an output signal -2 (in; supplies. This output signal is applied to a second fixed contact 44 of the changeover switch 40. The movable contact 46 of the changeover switch 40 (or its electronic equivalent) is connected to the steering controller 48.

The steering controller 48 influences the movement of the missile via an actuator 50, the movement of the missile and the pathology being symbolized again by a box 52. The line of sight rotation speed or " is measured by the seeker head 36 and converted into a corresponding signal op, with which the normal steering loop is closed.

The distance Rr of the missile from the target is detected by means of a proximity sensor 54, and a corresponding signal is applied to a threshold switch 56. The threshold switch 56 supplies an output signal Y- 0 for Rr>Rh; and an output signal K = Δ for R _T <Rri; - The L signal at the output of the threshold switch 56 moves the changeover switch 40 from the position shown, in which the contact 38 is connected to the input of the steering controller 48, to the other position , where the output of the threshold switch 42 is switched to the steering controller 48 via the contact 44.

If the distance Rt is smaller than the limit value Rn, then the switch 40 is thrown and the (»/ • output of the seeker head is switched off by the steering controller 48. Instead, the steering controller 48 receives the output signal 2 dpa if λ> λ _{( ;} is or -2 (ip _C " if λ <λ <·; is. The steering controller 48 then commands the maximum possible transverse acceleration of the missile in any case, with the sign corresponding to the sign of the squint angle. Rtc was selected in this way been that this results in the desired lead.

The arrangement described was based on the control of a missile to a point in front of a "Hot Spot" explained. However, the controls can be used in the same way to control the missile to steer to a point which is otherwise in front of the point detected by the seeker head. When the seeker head z. B. If a jet engine arranged at the tail of the aircraft is approved, then it is generally cheaper to use the To direct missile to a point lying in front of this engine by a certain amount, because there a hit is more effective.

4 liliiti Zui

Claims (1)

Patent claims: 1. Control device for missiles that by means of a homing head on aircraft, containing: a homing head and means for continuously aligning the homing head with one Target, Means for generating a line of sight change signal which is proportional to the rate of change in space of the line of sight running from the target seeker head to the target, a steering controller acted upon by the line of sight change signal,
a sensor for the "squint angle" between the missile's longitudinal axis and the optical axis of the homing head and a threshold switch, via which an output signal dependent on dpra squint angle can be switched to the steering controller, characterized in that (a) the threshold value switch (18, 42) supplies an output signal with a corresponding sign when a positive or below a negative threshold value (± λα) of the squint angle (λ) is exceeded, (b) a proximity sensor (24, 54) is provided which supplies a distance signal (Ri) in accordance with the spacing between the missile (FK) and the aircraft (FZ) , and (c) the output signal of the threshold switch (18, 42) and the Absiand. ^c Mgnal (Ri) are connected to a logic circuit (26, 22 or 56, 40) to form a lead signal which acts on the steering controller (12, 48) and which determines the trajectory of the missile at small distances (Ri) and large squint angles (λ) to a point offset in the direction of the aircraft speed vector (Vi) with respect to the target aimed at by the homing head (10, 36).